NACA M14 AIRFOIL (m14-il)
NACA M14 AIRFOIL - NACA/Munk M-14 airfoil
| Details | Dat file | Parser | |
|
(m14-il) NACA M14 AIRFOIL NACA/Munk M-14 airfoil Max thickness 8.1% at 30% chord. Max camber 3.9% at 30% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
NACA M14 AIRFOIL
17. 17.
0.0000000 0.0000000
0.0125000 0.0153100
0.0250000 0.0230300
0.0500000 0.0346500
0.0750000 0.0441800
0.1000000 0.0518000
0.1500000 0.0636500
0.2000000 0.0721000
0.3000000 0.0800000
0.4000000 0.0797000
0.5000000 0.0732000
0.6000000 0.0625000
0.7000000 0.0491000
0.8000000 0.0346000
0.9000000 0.0195000
0.9500000 0.0119500
1.0000000 0.0042000
0.0000000 0.0000000
0.0125000 -.0110900
0.0250000 -.0120700
0.0500000 -.0119500
0.0750000 -.0108200
0.1000000 -.0094000
0.1500000 -.0063500
0.2000000 -.0036000
0.3000000 -.0006000
0.4000000 -.0009000
0.5000000 -.0025000
0.6000000 -.0044000
0.7000000 -.0059000
0.8000000 -.0055000
0.9000000 -.0042000
0.9500000 -.0024500
1.0000000 0.0000000
|
No parser warnings |
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for NACA M14 AIRFOIL (m14-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| m14-il | 50,000 | 9 | 38.1 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m14-il | 50,000 | 5 | 38.8 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m14-il | 100,000 | 9 | 57.3 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m14-il | 100,000 | 5 | 57.4 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m14-il | 200,000 | 9 | 76.8 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m14-il | 200,000 | 5 | 75.4 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m14-il | 500,000 | 9 | 102.9 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m14-il | 500,000 | 5 | 98.1 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m14-il | 1,000,000 | 9 | 121.9 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m14-il | 1,000,000 | 5 | 113.7 at α=4° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||